The present invention relates generally to gas turbine engines, and, more specifically, to compressors therein.
A typical axial flow compressor in a gas turbine engine includes several rows of compressor rotor blades extending radially outwardly from corresponding rotor disks, with corresponding rows of stator vanes being disposed therewith. Adjacent rows of vanes and blades define a compressor stage, with each stage further compressing air in-turn in the axial downstream direction. The finally compressed air is then mixed with fuel in a combustor and ignited for generating hot combustion gases which flow through conventional turbines which extract energy therefrom for powering the compressor in operation, as well as providing useful work in propelling an aircraft, for example.
The stator vanes may either be variable in angular orientation or fixed, and extend radially inwardly from the stator casing. The casing is typically split along an axial, centerline splitline or split plane into upper and lower casing halves, with each casing half having axial flanges at the splitline for bolting together the two casing halves to form the fully annular casing. The casing may also be split in the radial direction into adjoining front and rear portions if desired for long, multi-stage axial compressors. Corresponding radial flanges are used for bolting together the front and rear portions.
During assembly, the stator vanes are firstly mounted into their respective casing halves. Conventional vane shrouds are then mounted to the radially inner ends of the vanes for many of the vane stages as desired in a particular design. The various rotor stages are separately assembled together using interconnecting annular spacers therebetween to collectively form a rotor or spool with several stages of rotor blades extending radially outwardly therefrom and spaced axially apart for receiving the corresponding rows of stator vanes. The rotor is then assembled into position in the preassembled casing lower half with the rotor blades being disposed between corresponding stages of the stator vanes. The preassembled casing upper half, including the remaining stages of stator vanes joined thereto, is then assembled over the casing lower half and pre-installed rotor, with the axial flanges being bolted together to complete the assembly of these components.
The assembly process requires sufficient axial clearance between the various vane shrouds and the forward and aft adjacent blade rows for allowing assembly without interference or obstruction. Since the vanes and shrouds are stationary components, and the rotor blades and disks rotate, the axial clearances between these components must be sufficient to prevent undesirable contact therebetween during operation of the engine, and during attendant thermal expansion and contraction of the components.
The annular regions defined between adjoining blade stages around corresponding spacer therebetween defines an annular cavity in which the vane shroud is contained. The cavity size is relatively large due to the axial clearance requirements and assembly considerations typically found in axial compressors. Extending radially outwardly from the interstage spacers are annular labyrinth seal teeth which cooperate with conventional annular seal members supported by the vane shroud for sealing airflow around the vane shrouds during operation. The labyrinth seals reduce backflow of the compressed air around the vane shrouds in the cavities due to the pressure gradient of the compressed air which has a higher pressure in each succeeding stage than that of the previous stage.
Although the labyrinth seals are affective for reducing this backflow leakage, relatively large cavities nevertheless are undesirable since they add to undesirable rotor windage pressure losses and secondary flows, and adversely affect compressor stall margin. Leakage of compressor air into the cavities during operation subjects the air to frictional acceleration by the rotor and secondary flows in the cavity therefrom.
Reducing the cavity size by reducing the axial clearances between the vanes and blades is undesirable and is limited by required minimum axial clearances for allowing both assembly of the components as well as thermal expansion and contraction of the components during operation. Increasing size of the vane shrouds to occupy more space in the cavities is undesirable due to the increased weight thereof, complexity, and cost, for example. In an aircraft gas turbine engine, weight is a major design factor, which should be maintained as low as possible for maximizing efficiency of operation.